A typical gas turbine engine, such as an axial flow gas turbine engine, includes a compressor section, a turbine section, and a combustor section. The combustor section is located between the compressor section and the turbine section. Air at atmospheric pressure is initially compressed by the compressor section, and the resulting compressed air is delivered to the combustor section. In the combustor section, heat is added to the compressed air leaving the compressor section by mixing fuel with the compressed air and by burning the resulting fuel/air mixture. The high temperature gas flow resulting from the combustion of the fuel/air mixture in the combustor section expands through the turbine section, and some of the energy of this high temperature gas flow is used to drive the turbine section in order to produce mechanical power.
A turbine section may have one or more stages, wherein each stage employs one row of stationary nozzle guide vanes and a corresponding row of blades. Each row of blades is mounted on a corresponding rotatable turbine wheel. A turbine wheel, for example, may be in the form of a disk. The nozzle guide vanes are aerodynamically designed to direct incoming gas from the combustor section onto the turbine blades to thereby aerodynamically transfer kinetic energy to the blades causing rotation of the turbine wheel.
In the past, the high temperature combustion gases entering the turbine section typically have had a gas entry temperature in the range of 850.degree. F. to at least 1200.degree. F. Since the efficiency and work output of the gas turbine engine are related to the gas entry temperature of the incoming high temperature combustion gases, there is a trend in gas turbine engine technology to increase the gas entry temperature. A consequence of increasing the gas entry temperature of the combustion gases in a gas turbine engine is that the materials of the nozzle guide vanes and blades must be chosen so that the nozzle guide vanes and blades can resist such increased gas entry temperatures.
Historically, nozzle guide vanes and blades have been made of metals, such as high temperature steels, and, more recently, such as nickel alloys. Even with these types of high temperature materials, it has been found necessary to provide internal cooling passages in order to prevent melting of these materials. Also, ceramic coatings can be applied to the nozzle guide vanes and blades to further enhance the heat resistance of such nozzle guide vanes and blades. In specialized applications, nozzle guide vanes and blades are being made entirely of ceramic, which resists even higher gas entry temperatures.
However, if the nozzle guide vanes and/or blades are made of ceramic, which has a different chemical composition, physical property, and coefficient of thermal expansion to that of their corresponding metal supporting structures, then undesirable stresses, a portion of which are thermal stresses, will result between the nozzle guide vanes and/or blades and their metal supporting structures when the gas turbine engine is operating. Such undesirable thermal stresses cannot effectively be contained by cooling.
Conventional joints between the blades and turbine wheels of a turbine section have typically used a fir tree, or a dove tail, root design. Historically, a dove tail root design has been used with a ceramic blade in order to attach the ceramic blade to a metal turbine wheel. A metal compliant layer is used between the highly stressed ceramic blade root and the metal turbine wheel in order to accommodate relative movement therebetween and any sliding friction which may occur as a result of this relative movement. Sliding friction between the ceramic blade and the metal turbine wheel creates a compact tensile stress on the ceramic blade that degrades the ceramic surface of the ceramic blade. This degradation in the ceramic surface of the ceramic blade occurs in a tensile stress zone of the blade root. Therefore, if a surface flaw is generated in the ceramic surface of critical size, the blade root fails catastrophically.
The present invention overcomes one or more of the problems as set forth above.